The present invention generally relates to power acquisition for spacecraft and, more particularly, to algorithms and methods of power acquisition for spacecraft in solar panel wing-deployed configuration.
Prior art spacecraft typically acquire the sun for power safety using sun sensor assemblies, for example, either wide field of view (WFOV) sun sensor or narrow field of view (NFOV) slit sun sensor. Prior art methods of sun, or power, acquisition using sun sensor assemblies often serve dual purposes of acquiring the sun to achieve an accurate spacecraft body-to-sun attitude knowledge and placing the solar wings normal, i.e., orthogonal, to the sun-line, i.e., a line from the spacecraft to the sun, to maximize solar power. A prior art method of power acquisition is disclosed in U.S. Pat. No. 5,255,879, issued Oct. 26, 1993, entitled xe2x80x9cThree Axes Stabilized Spacecraft and Method of Sun Acquisitionxe2x80x9d, and assigned to the assignee of the present invention. Methods, however, that use sun sensor assemblies for power acquisition can require expensive electronic hardware, such as buffer channel hardware, and can require precise maneuvering to maintain a perfect slew along a spacecraft axis that is not very close to one of the principal axes. In a spacecraft configuration with solar wing or wings deployed where accurate spacecraft body-to-sun attitude knowledge is not needed, but rather it is only needed to keep the spacecraft power safe and thermally safe, a simple and robust algorithm for power safety, which, for example, avoids the difficulties just described with sun sensor buffers and precise maneuvering, would be preferable.
For increased dependability and safety, modern spacecraft may rely on more than one spacecraft control processor (SCP) to control the functions and attitude of the spacecraft. A situation where the spacecraft switches control from one SCP to another is referred to as xe2x80x9ctogglexe2x80x9d. An example toggle situation occurs when a first spacecraft control processor, SCP1 and its connected sensors, is in active control of the spacecraft and a second spacecraft control processor, SCP2 and its connected sensors, is in standby, i.e., is powered off. SCP1 encounters some problems detected by a monitor and decides to turn off SCP1 and turn on SCP2. SCP2 then takes over control of the spacecraft.
With the advent of high power-dissipation payloads in spacecraft, comes a large radiator to maintain payload temperature below the high operating limit. There are various situations in the operation of the spacecraft when power to the spacecraft payload may be turned off. After a situation where the spacecraft toggles from one spacecraft control processor to another, power to the spacecraft payload may be turned off. In post-toggle, i.e., after a toggle has occurred, with the spacecraft in payload-off configuration, the radiator continues to dissipate a large amount of power and, thus, requires more power for the heater to maintain the spacecraft above survival temperature compared to spacecraft with lower power-dissipation payloads. If the payload remains on, the spacecraft needs even more power to survive and the situation is worse than payload off. A power acquisition algorithm, referred to as safe-hold algorithm, may be used post-toggle to ensure that the spacecraft achieves power, thermal, and momentum safety. Due, for example, to the increased post-toggle power requirements for the heater with high power-dissipation payloads, the previously used safe-hold algorithm is no longer adequate to ensure power, thermal, and momentum safety of the spacecraft under all conditions or scenarios.
For example, a previously used safe-hold algorithm for post-toggle safety rotates the spacecraft body along an axis in the xz plane (perpendicular to the wing axis) and places the solar wing in sun searching mode. In the worst case, referred to as xe2x80x9cpaddle wheel scenarioxe2x80x9d (or cases close to this scenario), the received wing power is a sinusoidal curve with theoretical average wing power of 2/xcfx80, which is approximately 64%. The paddle wheel scenario occurs when the sun line and the spacecraft spin axis are orthogonal to each other, resulting in the sun being in the xe2x80x9ckeyholexe2x80x9d, i.e., sun in wing axis or the body y-axis, twice per revolution. Due to bus shadowing effect and the fact that the wing will take more than 360 seconds to rotate 180 degrees to search for the sun every time the sun passes the keyhole position, the received wing power is lower than the theoretical value. The useful wing power is further limited by the battery charge rate so that some of the power is discarded because the battery cannot accept all received power for charging. Because the spacecraft requires power higher than this useful wing power for safety, the previously used safe-hold algorithm no longer renders power, thermal, and momentum safety for the spacecraft.
FIG. 1 shows a phase transition block diagram of a previously used power acquisition algorithm. Algorithm 100, shown in FIG. 1, uses wide field of view (WFOV) sun sensor assembly (SSA), and is applicable to both wing stowed or wing deployed spacecraft configuration. Algorithm 100 includes an initialization phase 102, null rate phase 104, pitch search phase 106, keyhole search phase 108, steer to null phase 110, sun hold phase 112, and fault hold phase 114. The arrows shown in FIG. 1 indicate the control logic, or flow of control, between phases of algorithm 100. For example, from initialization phase 102 control may pass either to null rate phase 104 under a xe2x80x9cnormalxe2x80x9d state of affairs, or to fault hold phase 114 under abnormal conditions, such as algorithm 100 xe2x80x9ctiming outxe2x80x9d before power acquisition has been achieved. As seen in FIG. 1, pitch search phase 106 may be followed by keyhole search phase 108 or may proceed directly to steer to null phase 110. As described above, algorithm 100 may leave the spacecraft in the paddle wheel scenario where useful wing power is inadequate to place the spacecraft in power, thermal, and momentum safety. Thus, algorithm 100 represents a complicated algorithm for power acquisition, which is not very robust, i.e., is prone to failure, or entering the fault hold phase 118 state, under many conditions.
As can be seen, there is a need for a simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft, and especially for post-toggle power, thermal, and momentum safety for spacecraft with high heater-power demand in payload-off configuration. There is also a need for an algorithm for spacecraft power acquisition for power, thermal, and momentum safety that can rely on direct power measurement from the solar wings rather than the indirect power measurement provided by the use of sun sensor assemblies such as WFOV sun sensors.
The present invention provides a simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft, and especially for post-toggle power, thermal, and momentum safety for spacecraft with high heater-power demand in payload-off configuration. The present invention also provides an algorithm for spacecraft power acquisition for power, thermal, and momentum safety that can rely on direct power measurement from the solar wings rather than the indirect power measurement provided by the use of sun sensor assemblies such as WFOV sun sensors.
In one aspect of the present invention, an algorithm for a spacecraft includes a wing sun search phase, an xz slew phase, and a safe hold phase. In the wing sun search phase, a solar wing current is monitored against a low current threshold while a solar wing sun search is performed, the xz slew phase is entered when the solar wing current stays below the low current threshold until a solar wing current persistence timer expires, and the safe hold phase is entered when the solar wing current does not stay below the low current threshold until the solar wing current persistence timer expires. In the xz slew phase, the spacecraft is slewed about an axis close to the xz plane while the solar wing current is continuously monitored against a high current threshold, a high current persistence timer is reset when the solar wing current is below the high current threshold, and the spacecraft is stopped slewing and the safe hold phase is entered when the solar wing current stays above the high current threshold until the high current persistence timer expires. In the safe hold phase, the spacecraft is slewed about an axis close to a y axis while the solar wing current is continuously monitored against the low current threshold, a low current persistence timer is reset when the solar wing current is above the low current threshold, and the spacecraft is stopped stewing and the xz slew phase is entered when the solar wing current stays below the low current threshold until the low current persistence timer expires.
In another aspect of the present invention, a method for spacecraft power acquisition includes performing steps included in each of a null rate phase, a wing sun search phase, an xz slew phase, and a safe hold phase. The null rate phase includes steps of starting a null rate persistence timer and a phase timer; resetting the null rate persistence timer if the spacecraft body rate differs from a null rate target by a threshold; commanding transition to a wing sun search phase when a difference of the spacecraft body rate and the null rate target stays below the threshold until the null rate persistence timer expires; and commanding transition to the wing sun search phase when the phase timer expires. The wing sun search phase includes steps of monitoring a solar wing current against a low current threshold while a solar wing sun search is performed; beginning to perform an xz slew phase when the solar wing current stays below the low current threshold until a solar wing current persistence timer expires; and beginning to perform a safe hold phase when the solar wing current does not stay below the low current threshold until the solar wing current persistence timer expires. The xz slew phase includes steps of slewing the spacecraft about an axis close to the xz plane of the spacecraft while continuously monitoring the solar wing current against a high current threshold; resetting a high current persistence timer when the solar wing current is below the high current threshold; and nulling slewing of the spacecraft and beginning to perform the safe hold phase when the solar wing current stays above the high current threshold until the high current persistence timer expires. The safe hold phase includes steps of slewing the spacecraft about an axis close to a y axis while continuously monitoring the solar wing current against the low current threshold; resetting a low current persistence timer when the solar wing current is above the low current threshold; and nulling slewing of the spacecraft and beginning to perform the xz slew phase when the solar wing current stays below the low current threshold until the low current persistence timer expires.
In still another aspect of the present invention, an algorithm for a spacecraft includes a null rate phase, a wing sun search phase, an xz slew phase, and a safe hold phase. In the null rate phase, a null rate persistence timer and a phase timer are started, the null rate persistence timer is reset if the spacecraft body rate differs from a null rate target by a threshold, transition to wing sun search phase is commanded when a difference of the spacecraft body rate and the null rate target stays below the threshold until the null rate persistence timer expires, and transition to wing sun search phase is commanded when the phase timer expires.
In the wing sun search phase, a wing sun search phase timer is started, a solar wing current is monitored against a low current threshold while a solar wing sun search is performed, the xz slew phase is entered when the solar wing current stays below the low current threshold until a solar wing current persistence timer expires, the safe hold phase is entered when the solar wing current does not stay below the low current threshold until the solar wing current persistence timer expires, and the safe hold phase is entered when the wing sun search phase timer expires after resetting the solar wing current persistence timer once.
In the xz slew phase, the spacecraft is slewed about an axis close to the xz plane while the solar wing current is continuously monitored against a high current threshold, a high current persistence timer is reset when the solar wing current is below the high current threshold, and the spacecraft is stopped slewing and the safe hold phase is entered when the solar wing current stays above the high current threshold until the high current persistence timer expires.
In the safe hold phase, the spacecraft is slewed about an axis close to a y axis while the solar wing current is continuously monitored against the low current threshold, a low current persistence timer, which is longer than the high current persistence timer, is reset when the solar wing current is above the low current threshold, and the spacecraft is stopped slewing and the xz slew phase is entered when the solar wing current stays below the low current threshold until the low current persistence timer expires.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.